Method and system for improved blade tip clearance in a gas turbine jet engine

ABSTRACT

A low or high pressure turbine case is machined on its outside surface to form circumferential notches. The notches coincide with the internal locations of labyrinth seals for the blades, or with “hot spots” that have been identified. A stiffener ring is shrunk with an interference fit into each notch through inducing temperature differentials between the ring and the case. The compressive circumferential force exerted by each ring prevents the low or high pressure turbine case from expanding as much as it would otherwise, thus improving blade tip clearance or counterbalancing the “hot spots”, stiffening the case, and improving case cooling. In an alternate embodiment a hydraulic nut may be used to push the ring in place and held by a locking nut. Alternatively, C-rings, or multiple segmented rings, may be coupled together by hydraulic, electrical, or other means and actuated by a controller to exert adjustable compressive circumferential force.

CROSS REFERENCE TO RELATED APPLICATION

This application claims the benefit of U.S. Provisional Application No.60/571,701, filed on May 17, 2004, titled “METHOD AND SYSTEM FORIMPROVED BLADE TIP CLEARANCE IN A GAS TURBINE JET ENGINE.”

FIELD OF THE INVENTION

This invention relates to gas turbine jet engines, and more particularlyto the high pressure turbine case and low pressure turbine case of gasturbine jet engines, and even more particularly to improving theclearance of the blade tips within the interior of the high and lowpressure turbine cases, to stiffening the high and low pressure turbinecases, and to cooling the high and low pressure turbine cases.

BACKGROUND OF THE INVENTION

Since the development of the gas turbine jet engine, blade tip clearancewithin the interior of the casing has been a challenging problem. Bladetip and inter-stage sealing have taken on a prominent role in enginedesign since the late 1960's. This is because the clearance between theblade tips and surrounding casing tends to vary due primarily to changesin thermal and mechanical loads on the rotating and stationarystructures. On today's largest land-based and aero turbine engines, thehigh pressure turbine case (“HPTC”) and low pressure turbine case(“LPTC”) have such large diameters that they are more susceptible toexpanding excessively and becoming out-of-round, exacerbating the bladetip clearance problem.

Reduced clearance in both the HPTC and the LPTC can provide dramaticreductions in specific fuel consumption (“SFC”), compressor stall marginand engine efficiency, as well as increased payload and mission rangecapabilities for aero engines. Improved clearance management candramatically improve engine service life for land-based engines andtime-on-wing (“TOW”) for aero engines. Deterioration of exhaust gastemperature (“EGT”) margin is the primary reason for aircraft engineremoval from service. The Federal Aviation Administration (“FAA”)certifies every aircraft engine with a certain EGT limit. EGT is used toindicate how well the HPTC is performing. Specifically, EGT is used toestimate the disk temperature within the HPTC. As components degrade andclearance between the blade tips and the seal on the interior of thecasing increase, the engine has to work harder (and therefore runshotter) to develop the same thrust. Once an engine reaches its EGTlimit, which is an indication that the high pressure turbine disk isreaching its upper temperature limit, the engine must be taken down formaintenance. Maintenance costs for major overhauls of today's largecommercial gas turbine jet engines can easily exceed one milliondollars.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic diagram of the overall structure of a typicalgas turbine jet engine.

FIG. 2 shows a sectional schematic diagram of a low pressure turbinecase of a typical gas turbine jet engine.

FIG. 3 shows a sectional schematic diagram of the low pressure turbinecase of FIG. 2 fitted with stiffener rings in an embodiment of themethod and system for improved blade tip clearance, case deformation,and cooling of the present invention.

FIG. 4 shows a sectional schematic diagram of Section A of the lowpressure turbine case of FIG. 3, showing the stiffener ring about to beseated in an embodiment of the method and system for improved blade tipclearance, case deformation, and cooling of the present invention.

FIG. 5 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring about to be seated inanother embodiment of the method and system for improved blade tipclearance, case deformation, and cooling of the present invention.

FIG. 6 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the method and system for improved blade tip clearance,case deformation, and cooling of the present invention.

FIG. 7 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the method and system for improved blade tip clearance,case deformation, and cooling of the present invention.

FIG. 8 shows the improvement in clearance under load in an embodiment ofthe method and system for improved clearance of the present invention.

FIGS. 9A, 9B, and 9C show sectional schematic diagrams of a section of alow pressure turbine case having the stiffener ring positioned on thelow pressure turbine case with a hydraulic nut and secured with alocking nut in another embodiment of the method and system for improvedblade tip clearance, case deformation, and cooling of the presentinvention.

FIG. 10 shows a schematic diagram of a low pressure turbine case havingstiffener rings actuated by hydraulic, electric, or other means inanother embodiment of the method and system for improved clearance ofthe present invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the Figures, in which like reference numerals and namesrefer to structurally and/or functionally similar elements thereof, FIG.1 shows a schematic diagram of the overall structure of a typical gasturbine jet engine. Referring now to FIG. 1, Gas Turbine Jet Engine 100has Fan 102 for air intake within Fan Frame 104. High PressureCompressor Rotor 106 and its attached blades and stators force air intoCombustor 108, increasing the pressure and temperature of the inlet air.High Pressure Turbine Rotor 110 and its accompanying blades and statorsare housed within High Pressure Turbine Case 112. Low Pressure TurbineRotor 114 and its accompanying blades and stators are housed within LowPressure Turbine Case 116. The turbine extracts the energy from thehigh-pressure, high-velocity gas flowing from Combustor 108 and istransferred to Low Pressure Turbine Shaft 118.

FIG. 2 shows a sectional schematic diagram of a low pressure turbinecase of a typical gas turbine jet engine. Referring now to FIG. 2,Centerline 202 runs through the center of Low Pressure Turbine Case 204(shown in cross-section). Rotor 206 (shown in cross-section) has Blade208 attached thereto. One skilled in the art will recognize that manymore blades and stators would normally be present within Low PressureTurbine Case 204. Only one Blade 208 is shown for simplicity.

Labyrinth seal designs vary by application. Sometimes the labyrinthseals are located on the blade tips, and sometimes they are located onthe inside diameter of the cases as shown in FIG. 2. Labyrinth Seals 210(shown in cross-section) line the inside diameter of Low PressureTurbine Case 204 forming a shroud around each rotating Blade 208,limiting the air that spills over the tips of Blades 208. The shape ofLabyrinth Seals 210 is designed to create air turbulence between thetips of each Blade 208 and the corresponding Labyrinth Seal 210. The airturbulence acts as a barrier to prevent air from escaping around thetips of Blades 208. Blade Tip Clearance 212, defined as the distancebetween the tip of Blade 208 and Labyrinth Seal 210, will vary over theoperating points of the engine. The mechanisms behind Blade TipClearance 212 variations come from the displacement or distortion ofboth static and rotating components of the engine due to a number ofloads on these components and expansion due to heat. Axis-symmetricclearance changes are due to uniform loading (centrifugal, thermal,internal pressure) on the stationary or rotating structures that createuniform radial displacement. Centrifugal and thermal loads areresponsible for the largest radial variations in Blade Tip Clearance212.

Wear mechanisms for Labyrinth Seal 210 can be generally categorized intothree major categories: rubbing (blade incursion), thermal fatigue, anderosion. Engine build clearances in both high pressure and low pressureturbine cases are chosen to limit the amount of blade rubbing. Studieshave shown that improved blade tip clearances in the high pressure andlow pressure turbine cases result in significant life cycle cost (“LCC”)reductions.

As a cold engine is started, a certain amount of Blade Tip Clearance 212exists between each Labyrinth Seal 210 and the tip of Blades 208. BladeTip Clearance 212 is rapidly diminished as the engine speed is increasedfor takeoff due to the centrifugal load on Rotor 206 as well as therapid heating of Blades 208, causing the rotating components to growradially outward. Meanwhile, Low Pressure Turbine Case 204 expands dueto heating but at a slower rate. This phenomenon can produce a minimumBlade Tip Clearance 212 “pinch point.” As Low Pressure Turbine Case 204expands due to heating after the pinch point, Blade Tip Clearance 212increases. Shortly after Low Pressure Turbine Case 204 expansion, Rotor206 begins to heat up (at a slower rate than Low Pressure Turbine Case204 due to its mass) and Blade Tip Clearance 212 narrows. As the engineapproaches the cruise condition, Low Pressure Turbine Case 204 and Rotor206 reach thermal equilibrium and Blade Tip Clearance 212 remainsrelatively constant.

There is tremendous benefit in narrowing Blade Tip Clearance 212 duringthe cruise condition. This is where the greatest reduction in SFC can begained (longest part of the flight profile). On the other hand, it isgreatly desirable to avoid rubbing. Minimal clearance must be maintainedat takeoff to ensure thrust generation as well as keeping EGT below itsestablished limit. Hence, it has been the goal of many control systemsto attempt to maintain a minimal Blade Tip Clearance 212 while avoidingrubbing over the entire flight profile.

Engine temperatures play a huge role in determining the operationalBlade Tip Clearances 212. Gas turbine performance, efficiency, and lifeare directly influenced by Blade Tip Clearances 212. Tighter Blade TipClearances 212 reduce air leakage over the tips of Blades 208. Thisincreases turbine efficiency and permits the engine to meet performanceand thrust goals with less fuel burn and lower rotor inlet temperatures.Because the turbine runs at lower temperatures, while producing the samework, hot section components would have increased cycle life. Theincreased cycle life of hot section components increases engine servicelife (TOW) by increasing the time between overhauls.

Engine SFC and EGT are directly related to HPTC blade tip clearance. Onestudy has shown that for every 0.001 inch increase in HPTC blade tipclearance, SFC increases approximately 0.1%, while EGT increases one °C. Therefore, a 0.010 inch HPTC blade tip clearance decrease wouldroughly produce a one % decrease in SFC and a ten ° C. decrease in EGT.Military engines generally show slightly greater HPTC blade tipclearance influence on SFC and EGT due to their higher operating speedsand temperatures over large commercial engines. Improvements of thismagnitude would produce huge savings in annual fuel and enginemaintenance costs amounting to over hundreds of millions of dollars peryear.

Reducing fuel consumption also reduces aero engine total emissions.Recent estimates indicate that Americans alone now fly 764 million tripsper year (2.85 airline trips per person). The energy used by commercialaircraft has nearly doubled over the last three decades. The increasedfuel consumption accounts for thirteen % of the total transportationsector emissions of carbon dioxide (CO₂). Modern aero engine emissionsare made up of over seventy-one % CO₂ with about twenty-eight % water(H₂O) and 0.3% nitrogen oxide (NO₂) along with trace amounts of carbonmonoxide (CO), sulfur dioxide (SO₂), etc. Air transport accounts for2.5% (600 million tons) of the world's CO₂ Production. Emissions fromland-based engines, primarily for power generation, contributes amountsin addition to these totals. Clearly a reduction in fuel burn willsignificantly reduce aero and land-based engine emissions.

Current large commercial engines have cycle lives (defined as the timebetween overhauls) that vary significantly, ranging typically between3,000 to 10,000 cycles. The cycle life is primarily determined by howlong the engine retains a positive EGT margin. New engines or newlyoverhauled engines are shipped with a certain cold build blade tipclearance which increases with time. As the engine operating clearancesincrease, the engine must work harder (hotter) to produce the same workand is therefore less efficient. This increase in operating temperature,particularly takeoff EGT, further promotes the degradation of hotsection components due to thermal fatigue. Retaining engine takeoff EGTmargin by maintaining tight blade tip clearances can dramaticallyincrease engine cycle life. This could also lead to huge savings inengine maintenance over a period of years due to the large overhaulcosts.

Previous attempts at blade tip clearance management can generally becategorized by two control schemes, active clearance control (“ACC”) andpassive clearance control (“PCC”). PCC is defined as any system thatsets the desired clearance at one operating point, namely the mostsevere transient condition (e.g., takeoff, re-burst, maneuver, etc.).ACC, on the other hand, is defined as any system that allows independentsetting of a desired blade tip clearance at more than one operatingpoint. The problem with PCC systems is that the minimum clearance, thepinch point, that the system must accommodate leaves an undesired largerclearance during the much longer, steady state portion of the flight(i.e., cruise).

Typical PCC systems include better matching of rotor and stator growththroughout the flight profile, the use of abradables to limit blade tipwear, the use of stiffer materials and machining techniques to limit orcreate distortion of static components to maintain or improve shroudroundness at extreme conditions, and the like. Engine manufacturersbegan using thermal ACC systems in the late 1970's and early 1980's.These systems utilized fan air to cool the support flanges of the HPTC,reducing the case and shroud diameters, and hence blade tip clearance,during cruise conditions.

All of the approaches described above have significant problemsassociated with them. Some are quite expensive, others achieve littleresults, especially during cruise where the greatest advantages aregained, or require actuation through the case due to the lack of currenthigh temperature actuator capabilities, which raise secondary sealingissues and added weight and mechanical complexity. However, none of theapproaches heretofore attempted matches the effectiveness of the presentinvention.

FIG. 3 shows a sectional schematic diagram of the low pressure turbinecase of FIG. 2 fitted with stiffener rings in an embodiment of themethod and system for improved blade tip clearance, case deformation,and cooling of the present invention. Referring now to FIG. 3, themethod and system of the present invention may be applied to existinggas turbine jet engines, or may be incorporated into the design andbuild of new gas turbine jet engines. The method and system of thepresent invention is applicable to the HPTC as well as the LPTC, and thedescription of the invention and figures in relation to the LPTC alsoapply equally to the HPTC and is not limited to the LPTC.

Notches 302, which may be of several different geometries as describedin detail below, are manufactured circumferentially, typically throughmachining, into the outside diameter of Low Pressure Turbine Case 204 tocoincide with one or more locations of the Labyrinth Seals 210. Inaddition to locations corresponding to one or more of the locations ofthe Labyrinth Seals 210, notches may be machined circumferentially inlocations corresponding to “hot spots” that have been identified in LowPressure Turbine Case 204 through computer modeling, through monitoringsurface temperatures, or through visual inspections for cracks when theengine is overhauled. For existing engines, Low Pressure Turbine Case204 is typically removed in order to repair cracks resulting from thethese “hot spots”. After such repairs, groves may then be appliedthrough a weld repair through machining. The external rings would thenbe shrink interference fit in the grooves.

Stiffener Rings 304 (shown in cross section) are then shrinkinterference fit into each Notch 302. Since Low Pressure Turbine Case204 is conical in shape, each Stiffener Ring 304 will have a differentdiameter. In each case, the inside diameter of each Stiffener Ring 304will be slightly less than the outside diameter of its correspondingNotch 302. Each Stiffener Ring 304 is heated, starting with the largestdiameter Stiffener Ring 304. Heating causes each Stiffener Ring 304 toexpand, increasing the inside diameter to a diameter that is greaterthan the outside diameter of its corresponding Notch 302. Oncepositioned in Notch 302, Stiffener Ring 304 is allowed to cool, whichshrinks with an interference fit into its corresponding Notch 302.

FIG. 4 shows a sectional schematic diagram of Section A of the lowpressure turbine case of FIG. 3, showing the stiffener ring about to beseated in an embodiment of the method and system for improved blade tipclearance, case deformation, and cooling of the present invention.Referring now to FIG. 4, Notch 302 is manufactured circumferentiallywith a reverse taper in one embodiment of the invention. Angle 402 forthe taper will vary from case to case, ranging from just greater than 0°for a cylindrical case to an appropriate degree that would depend uponthe specific geometry of a conical case. Stiffener Ring 304 is machinedcircumferentially on its inside diameter to match this same taper. Eventhough Stiffener Ring 304 is shrink interference fit onto Low PressureTurbine Case 204, the taper adds extra security so that Stiffener Ring304 will not slip axially on Low Pressure Turbine Case 204, which couldpossibly happen if Notch 302 was manufactured flat without the taper.When Stiffener Ring 304 has been heated it expands, giving rise to RingClearance 404, enabling Stiffener Ring 304 to be positioned as shownagainst Heel 406 of Notch 302. As Stiffener Ring 304 cools, it shrinksin diameter and seats itself circumferentially into Notch 302. Atambient temperature, due to the smaller inside diameter of StiffenerRing 304 to the outside diameter of Notch 302, a shrink with aninterference fit results, with compressive circumferential force beingapplied to Low Pressure Turbine Case 204 by Stiffener Ring 304, andtensile circumferential force is applied to Stiffener Ring 304 by LowPressure Turbine Case 204.

In one example, Low Pressure Turbine Case 204 may be fifty inches inoutside diameter at the portion where Blade 208 and Labyrinth Seal 210are located. Low Pressure Turbine Case 204 is made of nickel-based superalloy, such as Inconel 718, as is Stiffener Ring 304 through a forgingprocess. Super alloy Inconel 718 is a high-strength, complex alloy thatresists high temperatures and severe mechanical stress while exhibitinghigh surface stability, and is often used in gas turbine jet engines.Heating Stiffener Ring 304 to a calculated temperature will causeStiffener Ring 304 to expand, yielding an appropriate Ring Clearance 404when Low Pressure Turbine Case 204 is at ambient air temperature ofapproximately seventy ° F. Alternatively, Low Pressure Turbine Case 204may be cooled with liquid nitrogen or other means to a calculatedtemperature to cause Low Pressure Turbine Case 204 to shrink indiameter, yielding an appropriate Ring Clearance 404 when Stiffener Ring304 is at ambient air temperature of approximately seventy ° F. Or, anappropriate Ring Clearance 404 may be achieved through a combination ofcooling Low Pressure Turbine Case 204 and heating Stiffener Ring 304,each to various calculated temperatures. Increasing or decreasing theinside diameter of Stiffener Ring 304 will result in more or lesscompressive circumferential force and tensile stress as required for aparticular application, and within the stress limits of the materialthat Stiffener Ring 304 is made from.

In addition, the machining for Low Pressure Turbine Case 204 may be donein a first direction, such as radially, and the machining for StiffenerRing 304 may be done in a second direction, such as axially, which ismore or less perpendicular to the first direction. Since machiningleaves a spiral, or record, continuous groove on the machined surfaces,the grooves on each surface will align in a cross-hatch manner to eachother, increasing the frictional forces between the two surfaces andreducing the potential for spinning of Stiffener Ring 304 within Notch302. The plurality of grooves on Stiffener Ring 304, which is typicallymade of a nickel-base super alloy, are harder than the plurality ofgrooves on Notch 302 of Low Pressure Turbine Case 204, which istypically made of titanium, or in other low pressure turbine casings,possibly steel or aluminum. The nickel-base super alloy grooves willdent into the softer titanium, steel, or aluminum grooves.Alternatively, Stiffener Ring 304 could simply be spot welded in one ormore locations to Notch 302, or bolted to one or more flanges secured toNotch 302, to keep Stiffener Ring 304 from spinning in relation to Notch302. Machining in cross directions would not be needed in this case.

By thus positioning Stiffener Rings 304 in the manner described, BladeTip Clearance 212 is improved, especially during cruise operation of theengine. The compressive circumferential force applied by the StiffenerRings 304 prevent Low Pressure Turbine Case 204 from expanding due toheat as much as it would otherwise expand. Stiffener Rings 304 may bemade of the same material as Low Pressure Turbine Case 204, or may bemade of a different material with a lower coefficient of expansion,which would increase the compressive circumferential force applied overthat of a stiffener ring of the same material as the case as thetemperature rises.

Heat is mainly dissipated from the outside surface area of Low PressureTurbine Case 204 by convection. Another benefit to adding StiffenerRings 304 to Low Pressure Turbine Case 204 is that heat is dissipated ata greater rate because Stiffener Rings 304 act as cooling fins, whichresults in cooler operating temperatures within Low Pressure TurbineCase 204, also contributing to less expansion and smaller Blade TipClearance 212. Also, Stiffener Rings 304 help to maintain roundness ofLow Pressure Turbine Case 204.

FIG. 5 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring about to be seated inanother embodiment of the method and system for improved blade tipclearance, case deformation, and cooling of the present invention.Referring now to FIG. 5, Notch 502 is machined circumferentially with achevron shape in one embodiment of the invention. Angle 508 may vary byapplication. Stiffener Ring 504 is machined circumferentially on itsinside diameter to match this same chevron shape. Even though StiffenerRing 504 is shrink interference fit onto Low Pressure Turbine Case 204,the chevron shape adds extra security so that Stiffener Ring 304 willnot slip off of Low Pressure Turbine Case 204, which could possiblyhappen if Notch 502 was manufactured flat without the chevron shape.When Stiffener Ring 504 has been heated it expands, giving rise to RingClearance 404, enabling Stiffener Ring 504 to be positioned as shownagainst Heel 506 of Notch 502. As Stiffener Ring 504 cools, it shrinksin diameter and seats itself circumferentially into Notch 502. Atambient temperature, due to the smaller inside diameter of StiffenerRing 504 to the outside diameter of Notch 502, a shrink with aninterference fit results, with compressive circumferential force beingapplied to Low Pressure Turbine Case 204 by Stiffener Ring 504, andtensile circumferential force is applied to Stiffener Ring 504 by LowPressure Turbine Case 204.

FIG. 6 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the method and system for improved blade tip clearance,case deformation, and cooling of the present invention. Referring now toFIG. 6, for aero applications, where added weight to the engine is aconcern, Stiffener Ring 604 is manufactured to have a profile that, whenseated as shown in FIG. 6, is substantially flush with the outer surfaceof Low Pressure Turbine Case 204. Notch 302 with a reverse taper asshown in FIG. 4 is machined into Low Pressure Turbine Case 204. Inaddition, based on the engine to be designed or to be retrofitted, Notch302 may be machined deeper, and/or wider, and Stiffener Ring 604 givenadded depth, and/or width, in order to meet the compressive and tensilecircumferential stress requirements.

FIG. 7 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the method and system for improved blade tip clearance,case deformation, and cooling of the present invention. Referring now toFIG. 7, for aero applications, where added weight to the engine is aconcern, Stiffener Ring 704 is manufactured to have a profile that, whenseated as shown in FIG. 6, is substantially flush with the outer surfaceof Low Pressure Turbine Case 204. Notch 502 with a chevron shape asshown in FIG. 5 is machined into Low Pressure Turbine Case 204. Inaddition, based on the engine to be designed or to be retrofitted, Notch502 may be machined deeper and/or wider, and Stiffener Ring 704 givenadded depth, and/or width, in order to meet the compressive and tensilestress requirements.

One skilled in the art will recognize that, in addition to the reversetaper and chevron designs for the notch and stiffener ring as shown inFIGS. 4-7, various other designs may be utilized to accomplish the samegoals. For example, the notch may have one or more ridges and channels,angular or undulating, that will match up with one or more channels andridges, angular or undulating, on the inside surface of the stiffenerring. Or, the notch and stiffener ring may have an inverted chevronshape. Many other such shapes may be envisioned without departing fromthe scope of the present invention.

FIG. 8 shows the improvement in blade tip clearance under load in anembodiment of the method and system for improved clearance of thepresent invention. Referring now to FIG. 8, Stiffener Ring 304 as shownin FIG. 4 has been shrink interference fit onto Low Pressure TurbineCase 204, and the engine is now under load, such as during cruiseoperation. Labyrinth Seal 210 and Low Pressure Turbine Case 204 withInner Surface 802 and Outer Surface 804 are depicted with solid lines inthe positions they would be in without Stiffener Ring 304. Low PressureTurbine Case 204 would have expanded in diameter, and Labyrinth Seal 210would have moved away from Blade 208, giving rise to a wider Blade TipClearance 212. However, due to the compressive force exerted byStiffener Ring 304 on Low Pressure Turbine Case 204, Labyrinth Seal 210is in the position indicated in phantom as 210′, and Ring 304, InnerSurface 802 and Outer Surface 804 of Low Pressure Turbine Case 204 arein the positions indicated in phantom as 304′, 802′, and 804′, thusreducing Blade Tip Clearance 212′.

Thus, the present invention reduces the amount of expansion that wouldnormally occur due to heating in the LPTC and the HPTC, and consequentlyimproving blade tip clearance. As stated above, increased blade tipclearance accelerates the effects of low cycle fatigue and erosion dueto increased temperatures in the HPTC and LPTC, and degrades EGT marginand engine life. In general, for large gas turbine engines, blade tipclearance reductions on the order of 0.010 inch can produce decreases inSFC of one % and EGT of ten ° C. Improved blade tip clearance of thismagnitude can produce fuel and maintenance savings of over hundreds ofmillions of dollars per year. Reduced fuel burn will also reduceaircraft emissions, which currently account for thirteen % of the totalU.S. transportation sector emissions of CO₂. The present inventionreduces blade tip clearances at cruise condition to make a significantimpact on SFC and EGT margin and improving turbine efficiency. Theincreased outer surface area of the HPTC and LPTC due to the stiffenerrings in certain embodiments will increase cooling and result in lowerinternal temperatures which will lengthen the cycle life of the engine.Another result of the present invention is an increase in payload perengine due to the improvement in blade tip clearance. Additional poundsof freight may be transported per takeoff and landing. The presentinvention could easily replace more expensive passive clearance controloptions.

FIGS. 9A, 9B, and 9C show sectional schematic diagrams of a section of alow pressure turbine case having the stiffener ring positioned on thelow pressure turbine case with a hydraulic nut and secured with alocking nut in another embodiment of the method and system for improvedblade tip clearance, case deformation, and cooling of the presentinvention. Referring now to FIG. 9A, Stiffener Ring 904 is sized to fitwithout pressure in a location near an internal Blade 208 and LabyrinthSeal 210, or previously identified “hot spot”, and placed in positionthere. Next, a Hydraulic Nut 902 is threadably mounted to Low PressureTurbine Case 204. Hydraulic Nut 902 has Piston 906 which engages withStiffener Ring 904.

In FIG. 9B, Piston 906 has extended from Hydraulic Nut 902, pushingStiffener Ring 904 toward the larger diameter end of Low PressureTurbine Case 204, thus positioning Stiffener Ring 904 in the optimumlocation in relation to the internal Blade 208 and Labyrinth Seal 210and resulting in an interference fit. The amount that Piston 906 isextended by Hydraulic Nut 902 is calculated to produce a desiredcompressive circumferential force by Stiffener Ring 904.

In FIG. 9C, Hydraulic Nut 902 has been removed, and Locking Nut 908 hasbeen threadably attached in its place onto Low Pressure Turbine Case204. Retainer 910 of Locking Nut 908 engages with Stiffener Ring 904,thus securing Stiffener Ring 904 in place. This process is repeated foras many stages as required based upon turbine design. This embodiment ofthe invention may add excessive weight and would most likely be bestsuited for land based applications where weight is not of such concern.

FIG. 10 shows a schematic diagram of a low pressure turbine case havingstiffener rings actuated by hydraulic, electric, or other means inanother embodiment of the method and system for improved clearance ofthe present invention. Referring now to FIG. 10, Low Pressure TurbineCase 1000 has Stiffener C-Rings 1004 positioned at predeterminedlocations to coincide with blade/labyrinth seals and/or “hot spots”. Inthis embodiment of the invention, Stiffener C-Rings 1004 are not shrinkinterference fit onto Low Pressure Turbine Case 1000. A notch for eachStiffener C-Ring 1004 is still machined into Low Pressure Turbine Case1000, but the stiffener rings are c-rings rather than continuous rings.Each end of Stiffener C-Ring 1004 is linked to an Actuator Means 1002,which when actuated, pulls each end of Stiffener C-Ring 1004 together,exerting compressive force on Low Pressure Turbine Case 1000. The insidesurface of each Stiffener C-Ring 1004, or the notch surface, or both,may be coated with Teflon® or some other lubricating substance tofacilitate slippage when tightened.

Each Actuator Means 1002 is connected to Controller 1008 throughElectrical/Electronic Connections 1006. Controller 1008 receivestemperature readings from multiple temperature sensors located near eachStiffener C-Ring 1004 (not shown). It is also possible to derive theLPTC temperature from EGT temperature readings and use these readingsfor feedback to Controllers 1008. As the temperatures being monitoredthroughout Low Pressure Turbine Case 1000 rise, Controller 1008processes the temperature data and determines how much each of the endsof each Stiffener C-Ring 1004 need to be pulled together by eachActuator Means 1002 in order to exert the proper compressivecircumferential force on Low Pressure Turbine Case 1000 to eithermaintain an optimum blade tip clearance or counterbalance the “hotspot”.

In an alternate embodiment, instead of a c-ring, a chain-like multiplesegmented ring may be coupled together by Actuator Means 1002. Inanother embodiment of the invention, the stiffener rings may be made ofa strip of non-metallic material, such as Kevlar®. The inside surface ofthe Kevlar®, or the notch surface, or both may also be coated withTeflon® or some other lubricating substance to facilitate slippage whentightened.

Having described the present invention, it will be understood by thoseskilled in the art that many and widely differing embodiments andapplications of the invention will suggest themselves without departingfrom the scope of the present invention.

1. A method for improved blade tip clearance in a gas turbine jetengine, the method comprising the steps of: (a) machining at least onestiffener ring to fit without pressure near a predetermined location onan outer surface of a turbine case of the gas turbine jet engine; (b)mounting a hydraulic nut on said turbine case adjacent to said at leastone stiffener ring; (c) engaging a piston of said hydraulic nut withsaid at least one stiffener ring; (d) pushing said at least onestiffener ring by extending said piston of said hydraulic nut in adirection toward a larger diameter end of said turbine case resulting inan interference fit of said at least one stiffener ring into saidpredetermined location; (e) replacing said hydraulic nut with a lockingnut, said locking nut having a retainer; and (f) engaging said retainerof said locking nut with said at least one stiffener ring, wherein saidat least one stiffener ring is secured in said predetermined location insaid interference fit; wherein said at least one stiffener ringdissipates heat, applies compressive circumferential force to saidturbine case, and prevents said turbine case form going out-of-roundduring cruise operation of the gas turbine jet engine, improving bladetip clearance.
 2. The method according to claim 1 wherein pushing step(d) further comprises the step of: extending said piston of saidhydraulic nut by an amount calculated to produce a desired saidcompressive circumferential force on said turbine case by said at leastone stiffener ring at said predetermined location.
 3. The methodaccording to claim 1 wherein machining step (a) further comprises thestep of: machining said at least one stiffener ring from a nickel-basesuper alloy.
 4. The method according to claim 1 wherein machining step(a) further comprises the step of: machining said at least one stiffenerring from a material that is different from a material of said turbinecase, said material having a lower coefficient of expansion than saidmaterial of said turbine case.
 5. The method according to claim 1wherein machining step (a) further comprises the step of: machining saidat least one stiffener ring to fit at an outer surface of said turbinecase at a location coinciding with a labyrinth seal on an inner surfaceof said turbine case.
 6. The method according to claim 1 whereinmachining step (a) further comprises the step of: machining said atleast one stiffener ring to fit at an outer surface of said turbine caseat a location coinciding with a hot spot of said turbine case.
 7. Anapparatus for improving blade tip clearance in a gas turbine jet engine,the apparatus comprising: at least one stiffener ring machined to fitwithout pressure near a predetermined location on an outer surface of aturbine case of the gas turbine jet engine; a hydraulic nut mounted onsaid turbine case adjacent to said at least one stiffener ring; a pistonof said hydraulic nut which engages with said at least one stiffenerring and pushes said at least one stiffener ring by extending from saidhydraulic nut in a direction toward a larger diameter end of saidturbine case resulting in an interference fit of said at least onestiffener ring into said predetermined location; a locking nut forreplacing said hydraulic nut; and a retainer of said locking nut whichengages with said at least one stiffener ring and secures said at leastone stiffener ring in said predetermined location in said interferencefit; wherein said at least one stiffener ring dissipates heat, appliescompressive circumferential force to said turbine case, and preventssaid turbine case form going out-of-round during cruise operation of thegas turbine jet engine, improving blade tip clearance.
 8. The apparatusaccording to claim 7 wherein said piston of said hydraulic nut isextended by an amount calculated to produce a desired said compressivecircumferential force on said turbine case by said at least onestiffener ring at said predetermined location.
 9. The apparatusaccording to claim 7 wherein said at least one stiffener ring ismachined from a nickel-base super alloy.
 10. The apparatus according toclaim 7 wherein said at least one stiffener ring is machined from amaterial that is different from a material of said turbine case, saidmaterial having a lower coefficient of expansion than said material ofsaid turbine case.
 11. The apparatus according to claim 7 wherein saidpredetermined location for machining said at least one notchcircumferentially into said outer surface of said turbine case is at alocation coinciding with a labyrinth seal on an inner surface of saidturbine case.
 12. The apparatus according to claim 7 wherein saidpredetermined location for machining said at least one notchcircumferentially into said outer surface of said turbine case is at alocation coinciding with a hot spot of said turbine case.
 13. A methodfor improved blade tip clearance in a gas turbine jet engine, the methodcomprising the steps of: (a) machining at least one notchcircumferentially at a predetermined location into an outer surface of aturbine case of the gas turbine jet engine; (b) coating an insidesurface of a stiffener ring with a lubricating substance, said stiffenerring having a first end and a second end; (c) seating said stiffenerring in each said at least one notch; (d) linking said first end andsaid second end of said stiffener ring to an actuator; and (e) actuatingsaid actuator to pull said first and second ends of said stiffener ringtogether, wherein said coating facilitates slippage of said stiffenerring in said at least one notch when said actuator is actuated; whereinsaid stiffener ring dissipates heat, applies compressive circumferentialforce to said turbine case, and prevents said turbine case form goingout-of-round during cruise operation of the gas turbine jet engine,improving blade tip clearance.
 14. A method for improved blade tipclearance in a gas turbine jet engine, the method comprising the stepsof: (a) machining at least one notch circumferentially at apredetermined location into an outer surface of a turbine case of thegas turbine jet engine; (b) coating a surface of said at least one notchwith a lubricating substance; (c) seating a stiffener ring in each saidat least one notch, said stiffener ring having a first end and a secondend; (d) linking said first end and said second end of said stiffenerring to an actuator; and (e) actuating said actuator to pull said firstand second ends of said stiffener ring together, wherein said coatingfacilitates slippage of said stiffener ring in said at least one notchwhen said actuator is actuated; wherein said stiffener ring dissipatesheat, applies compressive circumferential force to said turbine case,and prevents said turbine case form going out-of-round during cruiseoperation of the gas turbine jet engine, improving blade tip clearance.15. An apparatus for improving blade tip clearance in a gas turbine jetengine, the apparatus comprising: at least one notch machinedcircumferentially into an outer surface of a turbine case of the gasturbine jet engine at a predetermined location; a stiffener ring seatedin each said at least one notch, said stiffener ring having a first endand a second end; a lubricating substance coated on an inside surface ofsaid stiffener ring; and an actuator, wherein said first and second endsare linked to said actuator and said actuator when actuated pulls saidfirst and second ends together, said lubricating substance facilitatingslippage of said stiffener ring in said notch when said actuator isactuated; wherein said stiffener ring dissipates heat, appliescompressive circumferential force to said turbine case, and preventssaid turbine case form going out-of-round during cruise operation of thegas turbine jet engine, improving blade tip clearance.
 16. An apparatusfor improving blade tip clearance in a gas turbine jet engine, theapparatus comprising: at least one notch machined circumferentially intoan outer surface of a turbine case of the gas turbine jet engine at apredetermined location; a stiffener ring seated in each said at leastone notch, said stiffener ring having a first end and a second end; alubricating substance coated on a surface of said notch; and anactuator, wherein said first and second ends are linked to said actuatorand said actuator when actuated pulls said first and second endstogether, said lubricating substance facilitating slippage of saidstiffener ring in said notch when said actuator is actuated; whereinsaid stiffener ring dissipates heat, applies compressive circumferentialforce to said turbine case, and prevents said turbine case form goingout-of-round during cruise operation of the gas turbine jet engine,improving blade tip clearance.